High density metal-containing propellants capable of maximum boost velocity



June 20. 1967 A. c. SCURLOCK ETAL. 3,326,732

HIGH DENSITY METAL-CONTAINING PROPELLANTS CAPABLE OF MAXIMUM BOOST VELOCITY Filed Oct. 28, 1960 '7 SheetsSheet 240 5 01/04/70 FAT 48 I NVE N TORS 4000 6. 600000 4 AZVrH 5 @Me'a "1 M11000 [[0 fizz WK/606D AGENT June 20, 1967 HIGH DENSiTY Filed Oct. 28, 1960 A C. SCURLOCK ETAL METALCONTAINING PROPELLANTS CAPABLE OF MAXIMUM BOOST VELOCITY '7 Sheets-Sheet f5 INVENTORS 490/ 6. 56061006,

A275 5 5/4/1551 Mime [5 4%;

AGENT June 20. 1967 A, c. SCURLOCK ETAL 3,326,732

HIGH DENSITY METAL-CONTAINING PROFELLANTS E OF MAXIMUM BOOST VELOCITY CAPABL '7 Sheets-Sheet 4 Filed Oct. 28, 1960 a w H are e/ 9000 04/5; 6 70 00 7/0/1/ PAT/0:

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HIGH DENSITY METAL-CONTAINING PROPELLANTS CAPABLE OF MAXIMUM BOOST VELOCITY Filed Oct. 28, 1960 7 Sheets-Sheet 5 M2 INVENTORS 420 C. Scaewog A2709 Z /F//15 q Mm [55/655 AGENT Jun 20. 1967 A. c. SCURLOCK ETAL 3,326,732

HIGH DENSITY METAL-CONTAINING PROPELLANTS CAPABLE OF MAXIMUM BOOST VELOCITY Filed Oct. 28, 1960 7 Sheets-Sheet 6 /96 [DAT/01V A /o.-

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INVENTORS AFC/l 6." ficaema/g A2754 flmaa 4424420 [[5 65 BY W United States Patent HIGH DENSITY METAL-CONTAINING PRO- PELLANTS CAPABLE OF MAXIMUM BOOST VELOCITY Arch C. Scurloclr, Arlington, Keith E. Rumhel, Falls Church, and Millard Lee Rice, Annandale, Va., assignors to Atlantic Research Corporation, Fairfax, Va., a corporation of Virginia Filed Oct. 28, 1960, Ser. No. 65,860 20 Claims. (Cl. 14919) This invention rel-ates to new propellent compositions. More specifically it relates to propellent compositions of high propulsive performance, which contain high density, metal fuel components.

The term propellant, as employed herein, refers to compositions comprising an organic fuel binder or matrix containing molecularly combined carbon and hydrogen, an oxidizer, and a finely-divided metal or metal hydride fuel. When introduced into the combustion chamber of a rocket motor and ignited such compositions generate high temperature gases, which vent out through the restricted nozzle of the motor at a high velocity to produce thrust.

The parameter generally accepted as indicative of the propulsive performance of a rocket propellant is its specific impulse, which is defined as the lbs. of thrust generated per lb. of propellant per second. Although this characteristic of the propelant is very important as a criterion for evaluating its performance, specific impulse is not the sole factor determining optimum performance in a specific application where the rocket system as a whole, including its inert parts, such as weight of the motor casing and payload, must be considered.

The usual objectives of rocket propulsion are the following:

(1) Maximization of altitude and distance attained (2) Minimization of time-to-target (3) Maximization of thrust Each of these objectives is accomplished by maximizing the boost velocity imparted to the rocket device by combustion of the propellant, the boost velocity being the velocity of the rocket at burn-out less velocity at launchmg.

We have found that the relationship between boost velocity and the characteristics of the propellant, obtained by applying Newtons second law of motion to the rocket, is defined, for an idealized rocket free of gravitational and drag effects and pressure thrust, by the following equation:

AU=I.. 1 g [1 where AU=boost velocity=velocity at burn out less launching velocity M =mass of inert parts including payload=mass of rocket less mass of propellant V =volume occupied of propellant l specific impulse of the propellant g =dimensional conversion factor,

32.17 (lb. mass) (ft.)/(lb. force) (sec.)

=density of propellant M V =mass-to-volume ratio From this expression of the relationships which determine boost velocity, it will be seen that two additional factors, the density of the propellant and the ratio of the mass of inert parts to the volume of propellant enclosed within the motor, become exceedingly important, par- 3,326,732 Patented June 20, 11967 ticularly at high M /V values, as in the case of boosters and JATOs. Under such conditions, a propellant having a lower specific impulse than that of another propellant can be more effective in terms of boost velocity if the density of the former is substantially higher. The specific impulse of the composition is, of course, an important factor since, obviously, as between two propellants of the same density, the one having the higher specific impulse will impart a higher boost velocity to a given rocket vehicle.

The effect of the specific impulse and density charcteristics of the propellant and the M,/ V ratio is graphically shown in FIGURE 1, which compares the performance of a number of propellent compositions relative to a reference propellant, on propulsion of an idealized rocket vehicle, i.e. one for which the effects of drag, gravity, and pressure thrust are neglected, exhausting to 14.7 p.s.i. from a combustion chamber pressure of 1000 p.s.i. It will be noted that the effect of propellant density though substantial at relatively low M /V ratios, becomes increasingly important at such ratios increase. In the formulations shown, at high M /V ratios, the high density Zr propellent composition imparts a boost velocity which is as much as 27.5% higher than that imparted by the reference propellant, despite the lower specific impulse of the former. The low density oxygen-hydrazine propellant, on the other hand, despite its exceedingly high specific impulse, drops very considerably in propulsive performance vat high M /V ratios.

From the foregoing discussion, it is evident that the boost velocity which is particular propellant can impart to the rocket vehicle in a specific application can be a more important practical yardstick than the specific impulse of the composition.

The object of this invention is to provide propellent compositions of high density, which are effective to impart maximized boost velocity to rocket vehicles or devices.

Other objects and advantages will become obvious from the following detailed descriptions and the drawings in which:

FIGURE 1 is a graph comparing the performance of several propellants of different specific impulse and density at different ratios of the mass of the inert parts of the rocket vehicle to volume of propellant.

FIGURES 2 through 7 are graphs comparing performance in terms of boost velocity and specific impulse of propellent formulations containing zirconium, zirconium hydride, thorium, uranium, and hafnium respectively.

Broadly speaking, the invention comprises propellent compositions consisting essentially of a finely divided solid, inorganic oxidizer and a finely-divided metal or metal hydride fuel selected from the group consisting of zirconium, uranium, hafnium, and thorium, and the hydrides of said metals, dispersed in an organic fuel matrix containing molecularly combined carbon and hydrogen, in which the oxidizer and fuel components are present in a particular ratio by weight in terms of the number of atomic equivalents, defined as follows:

as free metal; and v equals the valence of the metal, preferably the characteristic valence of the metal in its most Oxidation Ratio:

3 stable oxide, which, in the case of Zr, Hf, Th and U, equals 4. Thus in the case of the following reaction:

the Oxidation Ratio 0.5

component to its oxide. Oxygen molecularly bonded to a carbon atom in an organic fuel compound, as, for example, in the case of an ether, alcohol, aldehyde, ketone, ester, amide, etc., though not available for combustion, nevertheless results, by decomposition reaction, in the formation of CO, the desired propulsive gas product. In cases such as organic acids, where decomposition might normally result in the production of CO under the conditions defined above for the Oxidation Ratio, one of the oxygen atoms preferentially combines with the metal fuel component.

When the components are present in a ratio by weight such that the Oxidation Ratio, as defined above, is 0.5, which will hereinafter be termed an OMOx formulation, the oxygen available for combustion of the fuel is stoichiometrically sufficient to oxidize the carbon not already linked to oxygen in the organic fuel to CO and the metal to its oxide. Under such conditions, the molecularly combined hydrogen in the organic fuel compound and in the metal hydride, if the latter is used, after ignition of the propellent composition, forms free hydrogen gas, since the molecularly combined carbon and the metal react preferentially with orxygen relative to the hydrogen. The hydrogen gas evolved is heated to a high temperature because of the high exotherm of the oxidizing metal. This, and its low molecular weight, make it a very efiicient thrust-producing component in the combustion reaction product.

We have discovered that propellent formulations containing Zr, U, Hf, and Th, and their hydrides, as highly densifying, highly exothermic fuel components, impart the highest boost velocity to a rocket system of relatively high M V ratio when the ratio of oxygen to fuel (organic and metal) is at OMOx or closely approaches OMOx, namely where the Oxidation Ratio equals about 0.48 to 0.60, the range 0.48 to 0.55 generally being preferred, although such formulations are not optimum with regard to specific impulse. These metal fuels, in such formulations, have the additional, highly advantageous characteristic of forming highly stable refractory oxides, which do not decompose or vaporize to any substantial degree after formation, either under the high temperature, high pressure conditions prevailing in the combustion chamber or during venting of the combustion gases out of the nozzle. Dissociation and/or vaporization of the metal oxides are undesirable since such phenomena absorb substantial amounts of heat energy, thereby reducing thetemperature and pressure of the low molecular weight, thrust-producing combustion gases.

These findings are graphically illustrated by the data summarized on the curves in FIGURES 2 through 7. In all of these figures, boost velocity was determined at an M V ratio of 60 lbs/cu. ft., assuming shifting chemical equilibria. In FIGURES 2 and 3, it will be seen that Zr at OMOx stoichiometry, both in a polyvinyl chloride fuel matrix plasticized with dibutyl sebacate and in a nitrocellulose-nitro'glycerine, double-base matrix, with ammonium perchlorate as the solid oxidizer, gave the highest boost velocity, although, in both cases, these were not formulations having the highest specific impulse. FIG- URES 4, 5, 6 and 7 show similar results for zirconium hydride, thorium, uranium, and hafnium.

The following comparative summary of the results obtained in double-base formulations, shows the effect of the density of the metal fuel component on boost velocity and the importance of the density factor relative to specific impulse at a high M /V ratio (60 lb./cu. ft).

Metal OMOx Metal Weight, Density Percent As aforementioned, the organic fuel matrix can be any suitable organic compound or mixture of organic compounds which contains molecularly combined carbon and hydrogen, so that at OMOx stoichiometry it burns and/or decomposes to produce CO and free hydrogen gas. It can be inert, the term inert as used herein meaning a compound which requires an external source of oxygen, namely the solid, inorganic oxidier, for combustion. Illustrative of suitable organic matrix compositions are the various solid polymeric binders, such as polyether polysulfide, polyurethanes, butadiene-acrylic acid and -metha crylic acid copolymers cross-linked With an epoxy, alkyd polyesters, polyamides, cellulose esters, e.g. cellulose acetate, cellulose ethers, e.g. ethyl cellulose, polyvinyl chloride, asphalt, and the like. The oxygen linked to carbon in a variety of such inert fuel binders produces CO by decomposition reaction.

Many of the solid polymeric binders preferably include high-boiling, organic, liquid plasticiers to improve physical properties and processing of the propellent composition. Any of the numerous organic plasticizers known in the art and compatible with the propellent composition can be employed. Illustrative examples of suitable organic plasticizers include sebacates such as dibutyl sebacate and dioctyl sebacate; phthalates, such as dibutyl phthalate and dioctyl phthalate; adipates, such as dioctyl adipate; glycol esters of higher fatty acids, and the like.

The organic fuel matrix can also comprise an active organic compound, a mixture of such compounds, or a mixture of such a compound with an inert organic compound, such as an inert organic plasticizer, the term active compound as employed herein meaning a compound which contains molecularly combined oxygen available for combustion of other components of the molecule, such as carbon. Examples of active organic fuel compounds include those containing nitroso, nitro, nitrite, and nitrate radicals, such as cellulose nitrate and nitroglycerine.

The foregoing description has dealt mainly with solid propellent compositions in which the organic fuel binder is a solid. The invention can also be employed in semisolid, composite monopropellent systems. Such compositions are thixotropic, cohesive, shape-retentive compositions which can be extruded under moderate pressures into the combustion chamber of a rocket, where they form continuously advancing columns which burn on the exposed surface. In accordance with this invention, such plastic monopropellent compositions comprise a stable dispersion of a finely divided, insoluble oxidizer and the finely-divided heavy metal or heavy metal hydride in a continuous matrix of any suitable high-boiling organic liquid fuel containing molecularly combined carbon and hydrogen. Illustrative of suitable liquid fuels are hydrocarbons, such as triethyl benzene, liquid polyisobutylene, and the like; organic esters, such as dimethyl maleate, dibutyl oxalate, dibutyl phthalate, and nitroglycerine; alcohols, such as benzyl alcohol and triethylene glycol; ethers, such as methyl fl-naphthyl ether; and many others.

Any solid inorganic oxidizer can be employed which yields oxygen readily for combustion of the metal fuel component and the organic matrix, where the latter contains no oxygen or insutlicient oxygen for CO stoichiometry. Such oxidizers include the inorganic oxidizer salts, such as NH K, Na, and Li perchlorates and nitrates, metal peroxides, such as CaO BaO and Na O and the like, the salts being preferred.

Where the organic fuel matrix contains molecularly combined oxygen available in at least the stoichiometric amount required for oxidation and/ or decomposition of the molecularly combined carbon component to CO, no inorganic oxidizer need be provided for its combustion. If such combined oxygen is present in amounts greater than that required for stoichiometry, the amount in excess preferentially oxidizes the powdered metal component rather than the molecularly combined hydrogen and thus can replace a portion of the inorganic oxidizer which would normally be required to oxidize the metal.

The organic fuel matrix, whether it be inert or active, solid or liquid, as aforedescribed, must comprise at least about 20%, preferably at least 30%, by volume of the propellent composition. This is essential both to provide for generation of an adequate amount of low molecular weight combustion gases requisite for effective propulsion and for processing of a cohesive propellent composition having good physical properties.

The amounts of metal or metal hydride and solid, inorganic oxidizer employed must be such as to produce, with the particular organic matrix, an Oxidation Ratio within the specified range of about 0.48 to 0.60. This can readily be calculated by use of the equation given above for determination of this expression.

The following examples are illustrative of propellent formulations within the scope of the invention.

Example 1 The following components were thoroughly blended at room temperature:

Parts by weight Nitrocellulose (12.6% N) (spheres -12 microns in diameter) 14.12 Nitroglycerine 12.07 Dibutyl phthalate 4.02 Dibutyl sebacate 1.15 Zirconium powder 45.90 Ammonium perchlorate 22.10 2'nitrodiphenylamine 0.64

Example 2 The following components were thoroughly blended at room temperature:

Parts by weight Polyvinyl chloride 6.46 Dioctyl adipate 8.29 Zr powder 40.60 Ammonium perchlorate 44.40 Wetting agent, equal parts glyceryl monooleate,

pentaerythritol dioleate, dioctyl sodium sulfosuccinate 0.25

The slurry thus formed was heated to 350 F., at which temperature the PVC dissolved in the dioctyl adipate to form a rigid gel matrix containing dispersed therein the Zr and NH ClO The Oxidation Ratio of the formulation is O'MOx. Burning rate of this composition at 75 F. was 0.5 in./sec. at 1000 p.s.i.

Example 3 A propellent composition was prepare-d as in Example 2 except that the components were in a somewhat different ratio by weight as follows:

This composition is slightly oxygen-rich, the Oxidation Ratio being 0.578. Burning rate of the grain, static-fired in a motor, was 0.49 in./sec. at a motor chamber pressure of 950 p.s.i.

Example 4 A propellent composition was prepared as in Example 1 except that hafnium powder was substituted for Zr and the component proportions were varied. Ratio by weight of the binder was the same.

Binder: nitrocellulose, nitroglycerine, dibutyl ph-thalate, dibutyl sebacate, 2-nitrodiphenylamine 30.0 Hf 59.3 NH ClO 10 7 This composition is OMOx. Burning rate at 1000 psi. was 0.44 in./ sec.

Although this invention has been described with reference to illustrative embodiments thereof, it will be apparent to those skilled in the art that the principles of this invention can be embodied in other forms but within the scope of the claims.

We claim:

1. In a rocket propellent composition which burns to produce propulsive gases and which consists essentially of a finely-divided, solid, inorganic oxidizer containing combined oxygen which it yields readily for combustion of the fuel components of said composition, and a finelydivided, solid metallic fuel component selected from the group consisting of Zr, U, Hf, Th and the hydrides of said metals, dispersed in an organic fuel matrix containing molecularly combined carbon and hydrogen, said organic fuel matrix comprising at least about 20 percent by volume of said composition, the improvement in which said oxidizer, said metallic fuel component and said organic fuel component are present in amounts such that the following expression:

equals about 0.48 to 0.60 wherein 0 equals the total amount of combined oxygen in the composition; C equals the total amount of carbon; M equals the total amount of said metallic fuel component calculated as free metal, said O, C and M being expressed in terms of the number of atomic equivalents; and v equals the valence of said metal.

2. The propellent composition of claim 1 in which the organic fuel matrix comprises at least about 30% by volume of the composition.

3. The propellent composition of claim 2 in which the oxidizer is an inorganic oxidizer salt.

4. The propellent composition of claim 1 in which the metallic fuel component is Zr.

5. The propellent composition of claim 2 in which the metallic fuel component is Zr.

6. The propellent composition of claim 3 in which the metallic fuel component is Zr.

7. The propellent composition of claim 6 in which the organic fuel matrix comprises an organic polymer.

8. The propellent composition of claim '7 in which the organic polymer is plasticized with a high-boiling, organic liquid plasticizer.

9. The propellent composition of claim 8 in which the organic polymer is polyvinyl chloride.

10. The propellent composition of claim 6 in which the organic fuel matrix comprises nitrocellulose plasticized with nitroglycerine.

11. The propellent composition of claim 6 in which the oxidizer salt is ammonium perchlorate.

12. The propellent composition of claim 7 in which the oxidizer salt is ammonium perchlorate.

13. The propellent composition of claim 9 in which the oxidizer salt is ammonium perchlorate.

14. The propellent composition of claim 10 in which the oxidizer salt is ammonium perchlorate.

15. The propellent composition of claim 2 in which the expression:

equals about 0.48 to 0.55.

16. The propellent composition of claim 1 in which v equals 4.

17. The propellent composition of claim 2 in which v equals 4.

18. The propellent composition of claim 5 in which v equals 4.

19. The propellent composition of claim 15 in which v equals 4.

20. The propellent composition of claim 19 in which the metallic fuel component is Zr.

Chem. and Eng. News, July 27, 1959, pp. 22, 23.

Zaehringer, Solid Propellant Rockets-Second Stage, American Rocket Co., Box 1112, Wyandotte, Mich. (1958), pp. 214-216.

BENJAMIN R. PADGETT, Primary Examiner.

LEON D. ROSDOL, ROGER L. CAMPBELL, Examiners. 

1. IN A ROCKET PROPELLANT COMPOSITION WHICH BURNS TO PRODUCE PROPULSIVE GASES AND WHICH CONSISTS ESSENTIALLY OF A FINELY-DIVIDED, SOLID, INORGANIC OXIDIZER CONTAINING COMBINED OXYGEN WHICH IT YIELDS READILY FOR COMBUSTION OF THE FUEL COMPONENTS OF SAID COMPOSITION, AND A FINELYDIVIDED, SOLID METALLIC FUEL COMPONENT SELECTED FROM THE GROUP CONSISTING OF ZR, U, HF, TH AND THE HYDRIDES OF SAID METALS, DISPERSED IN AN ORGANIC FUEL MATRIX CONTAINING MOLECULARLY COMBINED CARBON AND HYDROGEN, SAID ORGANIC FUEL MATRIX COMPRISING AT LEAST ABOUT 20 PERCENT BY VOLUME OF SAID COMPOSITION, THE IMPROVEMENT IN WHICH SAID OXIDIZER, SAID METALLIC FUEL COMPONENT AND SAID ORGANIC FUEL COMPONENT ARE PRESENT IN AMOUNTS SUCH THAT THE FOLLOWING EXPRESSION: O/(O+C+M(V/2)) EQUALS ABOUT 0.48 TO 0.60 WHEREIN O EQUALS THE TOTAL AMOUNT OF COMBINED OXYGEN IN THE COMPOSITION; C EQUALS THE TOTAL AMOUNT OF CARBON; M EQUALS THE TOTAL AMOUNT OF SAID METALLIC FUEL COMPONENT CALCULATED AS FREE METAL, SAID O, C AND M BEING EXPRESSED IN TERMS OF THE NUMBER OF ATOMIC EQUIVALENTS; AND V EQUALS THE VALENCE OF SAID METAL. 